Cooling of gas turbine rotors



Oct. 20, 1953 F. D. BROWNHILL ET AL COOLING OF GAS TURBINE ROTORS 8 Sheets-Sheet 2 Filed 001;. 8, 1947 FIG. 2.

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CoMPmsssoR Oct. 20, 1953 F. D. BROWNHILL ETAL COOLING OF GAS TURBINE ROTORS v Filed 001'.- 8, 1947 8 Sheets-Sheet 3 Oct. 20, 19,53 F. D. BR OWNH ILL ET AL 2,656,147

COOLING OF GAS TURBINE ROTORS F iled Oct. 8, 1947 8 Sheets-Sheet 4 9, 1953 F. D. BROWNHILL ET AL 2,656,147

I COOLING OF GAS TURBINE ROTORS Filed Oct. 8, 1947 8 Sheets-Sheet 5 Inventors: yrmw 2 F. D. BROWNHILL ETAL 47 COOLING oF'cAs TURBINE ROTORS Oct. 20, 1953 8 Sheets-Sheet 7 Filed 001;. 8, 1947 1953 F. D. BROWNHILL ET AL 6,

COOLING 0F GAS TURBINE RoToR Patented Oct. 20, 1953 COOLING OF GAS TURBINE ROTORS Frank Denison Brownhill and Paul Heinz Walter Wolfl", Rugby, England, assignors to The English Electric Company Limited, London, England, a British company Application October 8, 1947, Serial No. 778,618 In Great Britain October 9, 1946 The invention relates to cooling of the rotors of multi-stage gas turbines and has the object to provide a gas turbine rotor construction in which the only surfaces in contact with the hot working fluid are those of the blades proper and the circumference of a drum-shaped portion intermediate adjacent rows of rotor blades. The rotor discs that carry the blades are intended to work at considerably lower temperatures and are required to be free from thermal stresses such as would occur in a disc which is heated at the rim but remains comparatively cold at the centre.

According to the invention, in a multi-stage gas turbine the main body of the rotor and the outer portion of the rotor, which comprises the turbine blades, a series of blade platforms at the inner extremity of the blade profiles and a series of independent platforms located between the blade platforms of adjacent blade rings at the same diameter as the said blade platforms enclose between them a passage for a continuous and undiluted stream of cooling air forming an insulating wall between said main body and outer portion of the rotor, the only connection through the insulating stream of air being the lower portions of the blades which, between the said blade platform and the serrated roots by means of which they are secured into the body of the rotor, are reduced in cross-sectional area and made of sufiicient length to offer a substantial resistance to the flow of heat between the said outer hot portion of the rotor and the said main body, and the only connection between the said independent platforms and the said body of the rotor being through locating arrangements of relatively small area capable of offering considerable resistance to the flow of heat.

The invention is applicable both to axially inserted blades and to circumferential fixtures, both types having preferably serrated fixings of the so-called fir tree type. A reduced neck is formed between the blade root platform and the serrated fir tree stub so as to provide a narrow neck for the fiow of heat and at the same time a recess through which the cooling medium can flow. The cooling medium is preferably a small percentage of air tapped oif from the turbine compressor and emerges into the bladed zone of the turbine on the leading edge side of the rotor blades of each stage so as to form a comparatively cool boundary layer on the blade root before mixing with the combustion gases.

- In another form of embodiment high pressure cooling air is introduced into the insulating space 12 Claims. (01. 253-3915) i at one of the stages at the low pressure end of the rotor only, preferably the lowest stage, and then passes along substantially the whole length of the rotor in the insulating space through passages formed between adjacent blade stubs and between the reduced area portions of the blade root, these passages being shaped and proportioned so as to ensure a, high degree of heat abstraction from the blade roots and from the main body of the rotor. After passing along substantially the whole length of the rotor the cooling air isthen discharged into the main work ing fluid on the upstream side of one of the early stages of rotor blades, preferably the first stage.

This embodiment permits the economical use of larger total quantities of cooling air since the cooling air is returned to do useful work in the turbine at a pressure equal to that of the main working fluid after passing through the first stage nozzles. This together with the fact that the whole of the cooling air passes through each row of blade roots permits the use of compara tively large passages between them and thu min; imizes the danger of the passages becoming choked by foreign matter and at the same time promotes more efiicient cooling due to theincreased surface area.

V A further advantage of this embodiment is that heat abstracted by the cooling air from the low pressure portions of the turbine is returned to the cycle at a higher pressure and therefore with higher availability for conversion into work thus, making the cycle to some extent regenerative and increasing its efiiciency.

In order to be better understood and readily: carried into effect, the invention is illustrated by way of example in the accompanying drawings of which:

Fig. 1 is a side elevation in section of part of three consecutive rotor discs having axially inserted blades with series flow cooling means ar-.. ranged between adjacent discs, 1

Fig. 2 is part 'of a front elevation of a rotor disc and intermediate pieces corresponding to Fig. 1, g Fig. 3 is a side elevation in section of part of two consecutive rotor discs having axially inserted blades with parallel flow cooling means'ar.

Fig. 2,

nular space between themselves and the flanges a and 20b of the adjacent discs 2|, 2|. end of that bridge portion 3'! on the side of the The radial wall 32 forms a lip that leaves a comparatively narrow gap 36 between its end and the leading edge of the platform 2% of the adjacent blade 22. The other end of the bridge plate 31 has a radially inward pointing extension serrated-with the same fir tree profile as the blade attachment stubs 22a and of an axial dimension slightly smaller than the aforesaid circumferential groove 39b in the rotor disc. A rib 34 joins the platform 31, the serrated extension 35 and the wall portion 32. Adjacent platform sections 21 may be joined either by means of a spigotted joint 2162 (Fig. 3d) or by means of an inserted metal strip 2le (Fig. 36) forming a seal between the -annular space 28 and the stream of combustion gases passing the blades 22, 22', which will remain substantially airtight and will at the same time permit expansion of the said platform sections.

Before the assembly of the blades with the rotor the aforesaid separate intermediate pieces 21 are dropped with their serrated radial extensions 35 into the said circumferential groove 39a of the disc and then shifted in an axial direction until the projection 33 of the radial wall portion 32 on the other side of the bridge piece -31 engages the recess 26 between the flanges lel streams in a direction opposite to the flow of the combustion gases. These streams mix in another annular cavity 3| formed between the rim of each disc and a ring-shaped wall formed by the wall portions 32 of all of the separate intermediate pieces 21. The cooling medium emerges from this second annular cavity 3| through a circumferential gap 36 between the aforesaid lip :portions 39 of the intermediate pieces 2? and the :leading edges of the blade platforms 22b so as .to form a comparatively cool boundary layer covering the blade platforms before mixing with the hot combustion gases.

, Referring now to Figs. 4 to 4c showing circumferential fixtures, the intermediate pieces 41, bridging the space between two adjacent discs 4|, M respectively, are preferably oblique, i. e.

approximately of parallelogram shape in plan view. These pieces di are again secured by the engagement of circumferential recesses with circumferential flanges 45, a, 4552) respectively on .oppositefaces of adjacent discs. The last sec- 'tion lla (Fig. 4, right hand side) to be inserted is again of an axial dimension which is shorter than the distance between adjacent discs M, 31 so that it can be inserted radially and shifted axially to engagement with the one 4| of the adjacent discs, the remaining gap being closed by the radial insertion of a recessed piece 55 which can also be dropped in radially and shifted axially until it engages the other adjacent disc il", the remaining gap between these two por- :tions being filled by a wedge-shaped member iii, the two portions 46a, 56 and the said wedge- 6 shaped member 5i being eventually welded toether at 52.

The path of air (Figs. 4a and 4b) in this embodiment is preferably radially outward in the cavity between two adjacent discs 4 4 and 4|, 4| respectively, then at 46 in an axial direction between two adjacent intermediate pieces 41 into a circumferential space Ma left between the platform 4212 of the blades and the outer circumference of the rim of the disc. The air flows then radially inward through a recess left in the radial faces of adjacent blade attachment stubs 42a until it reaches another circumferential channel left at 43 between the inner end of the blade attachment stubs 42a. and the bottom of the circumferential groove lla left in the rim of the disc. From there the cooling air moves again radially outward in a recess left between the next pair of adjacent blade attachment stubs 420, into a circumferential channel 44b left between the outer circumference of the rim of the disc and the platform 42b of the blades, and from there in a substantially radial direction through a hole 420 arranged in the leading edge side of the blade root platform 321) so that the emerging air again forms a boundary layer of comparatively low temperature along the platforms of the rotor blades before mixing with the combustion gases.

Referring now to Fig. 5 the adjacent rotor discs 5| and 5| are again provided with spigotted flanges 56a, 591), respectively, which after assembly are welded with one another by the seam 59. The space thus enclosed between the extreme low pressure stage disc 51 and the adjacent disc 5| is connected with the source of cooling medium, preferably unburnt air of a pressure at least equal to the highest pressure occurring in the gas turbine cycle. A flange 7| is provided on the downstream side of disc 5| and is equipped with an inward axial ledge 12. Pieces in similar in shape to the intermediate platform pieces 2'! of Figs. 3 and 321 but generally shorter in the axial dimension are adapted to be inserted radially with their serrated inward extension 75 into the circumferential recess 59a of the disc 5| and shifted axially into the position shown in J'ig. 5 in which they engage with their extensions 73 the ledge 72 of the flange 1|. In this position they are held by the axial insertion from the left of the blades 52 having correspondingly serrated roots. The pieces 19 enclose between themselves and the disc 5| an air space 69 communicating with the space between the discs 5| and 5| by ducts 68.

Channels 53 and 54 are formed under the inner ends and under the platforms 52b of the blades, corresponding to the channels'f's, 24, respectively, of the embodiment of Fig. 3. The leading edges 52d of blades 52 have a spigotted joint with the flange 5!] of the intermediate platform pieces 51 which correspond otherwise substantially to the intermediate platform pieces 21 of Figs. 3 and 3b, in that they engage with their extensions 63 the grooves 56 between the circumferential projections 55, 55a of the discs 5|, 5| etc. The radial wall portions 62 of the intermediate platform pieces 5'! have wide openings 52a.

As will be seen from Fig. 5, this arrangement provides a series flow of the cooling medium from space 59 at the downstream side of the lowest stage through the channels 53, 5d holes 62a under the platforms 61 of the intermediate pieces 51 through the channels 53, 54 of the next higher stage, and so on, until the cooling medium is lasso, 147

channels 153, 5.4 :of f'liheli highest zpressureustage ihomever ercorresponding. otherwise ;tov zthe end zrpieces l0: atithe; low; ressure. end.

zlnrallithe .various embodiments::referred: to *hereabove, severe temperatureegradients inthe rroton can be -avoided'by applyingrcooling medhim ratzthee blade; roots.

Z'I'hustthe.rotorisrkeptrcool enoughtto minimise -creep in thezblade-attachment stubs, even when msmgwcomparatively low grade material L'for :the motor. Creep in the :blade .rootsthemselves 'is ssalsoz-rkept" to (:2, :min-imum. By controlling :the irpresslrre. of therxcooling: medium, any: inward I flow not the hot gasesis. prevented. As. stated aboveca aboundary-layerof cool. airrover'. the outside ofthe blades): and of "the: intermediate platforms; is set by the escaping; cooling; medium.

.The arim of. the :rotor discs ;of:.. any ".of the serfrhodiments xm-ay be rdivided into. segments by .YfidifilslSIOtS 1(notrshown) :in a way known in 2itselt,.:in order to.:allow anindependent thermal .7

.iexpansionofithat portionrwithout setting upondniy highicomp-ressivestresses.

' iiByitheaforesaid improvements .quick starting and changes of load and temperature ofthegas :turbine are: permissible.

:ZIhe 'construction :described: is "applicable also toztaperech rotors. Solid: forged. orbuilt up. rotors .maythe: used=and,: aseshown in the. embodiments described, :with .reierence to v:Figs. i3 and 15 the welding seamscof .built up; rotorszican be kept out =otrcontactnvith therstreami;of:comhustion gases, :a design which; isapplicable also sto other emi-hodim-ents.

The-method or securing the intermediate pieces adjacent blade rows.

What- We claim-asour' invention-and "desire to secure by' Letters Patentv is:

1. In "a multistage gas turbinea rotor coin- :prisingseveral rotor discsintegral with one another and each'coinprising a mainbody, proflled projections. of a cross sectional area restricted ascompared with that of'the said main :body extending from-its circumference, and an outer ;.portion comprising a. rowof blades. each blade-having an airfoileportion, a platform portion'and a root section of a crosssectional' area comparatively restricted adjacent. the said.platform portion and comparatively largerin. a portion profiled complementary to .and. engaging with the saidproflled projectionsofithe main "body; and separate platforms arranged, between the platform portions of adjacent rows of blades -:at" the same diameter asthe said platform portions, complementary coolant conducting .pas-

sagewaysformed in said profiled projections and said Tootsections'the-said root sections, plat- "form *portions and separate platforms forming enclosed passages between "them "and 'the said main body, means connecting said passages and passageways for conducting acontinuous stream of cooling air forming a heat insulating layer between the-said-main body andthe said outer "portiom the said root sections-forming the only integral connection between thesaid airfoil porad'apted' to be axially I inserted into :theesaid recessed and serrated projectionya circumferential :groove in the said projectionand an :opposi'te groove in each. blade root, theseegrooves' forming together a circumferential ductgthe said intermediate platforms having equallyserrated projections pointing inwardsof =an axialedimension I shorter than the. said a groove -.in *the rotor, radially inserted-into f the said: groove and axially moved into engagement with the: serrated 11'6- cesses of the rotor. and held-.therebythe serrated :bla'd'e roots, the .total axialilengthiof :the said serrated projections of thenintermediate :platforms plus the serrated blade rootsi-being .=substantially equal to 'theaXiaILIengthaof-"the F581?- rated recesses in the said projection: of. the:

rotor body.

3'. "In a multistage gas turbine asiset claim 1 -'cy1in'drical "flanges .on 517119 :si'des of a the said discs, adjacent flanges bemg-weldedstozgethen'the said intermediate:pIatformsrseparat ing the said welded flanges from the streamtof hot workingfluidipas'sing through the said-trotor :blades.

14. ?In :a :multistage gas turbine :a motor. :as claimed in 'claim 11 wherein the 1 said separate latforms' are anchoredini thesaidprojections of the main bodyz'andj form with; the adjacent rotor discs substantially closed i'drums exposed iinside to :the cooling; airflowing through -:thezrsaid ;pas-

sages.

5. Inxa multistage 'gas.turbineaaseset forthsin claim 14 317 least one. of. the said intermediateplat- .formxpiecescomprising two parts.:of-. a'.total: axial dimension. smaller than the 'distancebetween adiacent 'dlSCS'.S0"8.-S 'to beinserted iradially and :then shifted 1 axially into engagement with the .respective "adjacent disc, and wedge rmembers :drivenrradially between xthe-zsaidi partsro'f platformpieces and subsequently .welded .to timer-said parts.

6.. In: a multistage gas turbinezas; set. forth: in ciaim 4::meanstfor .circulatingca gaseousaicool'ing zmediumisuchr as compressed air-sthroughnthe said "passages and past: the. inside of the said intermediate platforms.

'i. .Inamultistage-gasxturbine aszset forth: in

:claim .6, the said tblades .tbeing iinserted axially into the said; prOjECti'OIillO'f the 'main'. rotortbody, "andthe said meansiforrcirculating thezcooling :medium directing the same; firstly axially; the

direction. of the main working :fluid -stream through. the said passages :under the inner ends of the blade roots of arotor. stage, then deflecting thesaid :cooling medium rradially. outwards, and

.directing the sai'd'cooli-ng' medium in a direction opposed to' that of the main working fluid stream through the said passages underthe'blade platin: the directionyiopposite :to: that :of the main working fluid; stream: both: through.' the saidma'ssages under theinnerblade root endsof a'rrotor stage and through theasaid passages .underith'e blade: platforms of; the: saidrotorstage, and eventually discharging-the:coolingmediumxatthe upstream side of the. rotor-blades voffthezsaid rotor s age.

9. In a multistage gas turbine as set forth in claim 6 cylindrical flanges on the said discs forming inner drums the outside radius of which is less than the radius defining the bottoms'of the recesses in the said outer portions, at least one of the said flanges being perforated by holes adapted to admit metered quantities of cooling medium from the space within the said cylindrical flanges to the said insulating spaces arranged in separate groups of the outer rotor portion, each group comprising at least one row of the said blade roots, blade platforms and intermediate platforms and'including the said passages for the cooling medium, the said cooling medium being discharged thereafter into the main Working fluid stream passing the blades proper of the gas turbine.

10. In a multistage gas turbine as set forth in claim 9 means for connecting the said insulating pressure stage, and the said cooling medium being discharged into the said main working fluid stream at the upstream side of the rotor blades of a higher pressure stage.

11. In a multistage gas turbine as set forth in claim 10 the said means for the supply of coolingmedium being connected to the insulating space on the downstream side of the lowest pressure stage of the rotor, the said means passing the said cooling medium consecutively through the 10 insulating spaces and passages 01' all stages of the said rotor, and eventually discharging it into the said main working fluid stream at the upstream side of the rotor blades of the highest pressure stage.

12. In a multistage gas turbine as set forth in claim 10 a hollow central portion of the rotor in supply connection on the one hand with a source of coolin medium, and on the other hand with the said insulating space of at least one rotor stage.

FRANK DENISON BROWNHILL. PAUL I-IEINZ WALTER WOLFF.

References Cited in the file of this patent UNITED STATES PATENTS Number Name Date 2,141,401 Martinka Dec. 27, 1938 2,149,510 Darrieus Mar. 7, 1939 2,213,940 Jendrassik Sept. 3, 1940 2,241,782 Jendrassik May 13, 1941 2,440,069 Bloomberg Apr. 20, 1948 2,452,782 McLeod et a1. Nov. 2, 1948 2,489,683 Stalker Nov. 29, 1949 FOREIGN PATENTS Number Country Date 452,412 Great Britain Aug. 24, 1936 543,985 Great Britain Mar. 23, 1942 578,191 Great Britain June 19, 1946 579.316 Great Britain July 31, 1946 

